Variable area turbine vane arrangement

ABSTRACT

A turbine section of a gas turbine engine includes a ring of turbine nozzle segments each having paired turbine vanes. Turbine throat area is modulated by rotating each rotational turbine vanes about an axis of rotation which is located such that rotation changes the turbine throat area concurrently between one rotational stator vane and two adjacent fixed turbine vanes. Each paired turbine vane doublet includes at least one rotational turbine vane between two fixed turbine vanes.

BACKGROUND OF THE INVENTION

The present invention relates to a gas turbine engine turbine section, and more particularly to a ring of paired turbine vanes doublets in which one vane of each doublet rotates to modulate turbine throat area.

The core engine of a gas turbine engine typically includes a multistage axial compressor which provides compressed air to a combustor wherein it is mixed with fuel and ignited for generating hot combustion gas which flows downstream through a high pressure turbine nozzle and in turn through one or more stages of turbine rotor blades. The high pressure turbine blades are joined to a rotor disk which is joined to the compressor by a corresponding drive shaft, with the turbine blades extracting energy for powering the compressor during operation. In a two spool engine, a second shaft joins a fan upstream of the compressor to a low pressure turbine disposed downstream from the high pressure turbine.

Typical turbine nozzles, such as high pressure and low pressure turbine nozzles, have fixed vane configurations and fixed turbine nozzle throat areas. Variable cycle engines are being developed to maximize performance and efficiency over subsonic and supersonic flight conditions. Some engines provide variability in compressor turbine vanes by mounting each vane on a radial spindle and collectively rotating each row of compressor vanes using an annular unison ring attached to corresponding lever arms joined to each of the spindles. Each compressor vane rotates about a radial axis, with suitable hub and tip clearances which permit rotation of the vanes.

Although it would be desirable to obtain variable flow through turbine nozzles by adjusting the throat areas thereof, previous attempts thereat have proved difficult because of severe operating environment of the turbine nozzles. The severe temperature environment of the turbine nozzle typically requires suitable cooling of the individual vanes, with differential temperature gradients. Nozzle vanes are also subject to substantial aerodynamic loads from the combustion gas during operation. Furthermore, adjustable turbine nozzle vanes may reduce the structural integrity and durability of the nozzle segments in view of the increased degree of freedom therebetween.

Accordingly, it is desirable to provide a variable area turbine nozzle having a relatively uncomplicated rotation, support and sealing structure to provide variable nozzle throat area capability yet minimize turbine pressure loss, leakage, expense and weight.

SUMMARY OF THE INVENTION

The turbine section of a gas turbine engine according to the present invention provides a ring of turbine nozzle segments each having paired turbine vanes (Doublets). Turbine throat area is modulated by rotating one of the doublet vanes about an axis located towards the rotatable turbine vane trailing edge. The vane adjacent each rotational turbine vane is a fixed turbine vane which provides a rigid structure to support the rotational turbine vanes which thereby form only a portion of the turbine section. The center of rotation of the rotational turbine vane is located such that rotation of the rotational turbine vane changes the turbine throat area concurrently between the rotational stator vane and both adjacent fixed turbine vanes.

The present invention therefore provides a variable area turbine nozzle having relatively uncomplicated rotation, support and sealing structure to provide variable nozzle throat area capability yet minimize turbine pressure loss, leakage, expense and weight.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently disclosed embodiment. The drawings that accompany the detailed description can be briefly described as follows:

FIG. 1 is a general perspective view an exemplary gas turbine engine embodiment for use with the present invention;

FIG. 2 is an expanded view of a vane portion of one turbine stage within a turbine section of the gas turbine engine, the vane portion formed from a multiple of turbine nozzle segments;

FIG. 3 is an expanded partial phantom view of one variable turbine nozzle segment; and

FIG. 4 is a top schematic representation of the throat change performed by the turbine section according to the present invention.

DETAILED DESCRIPTION OF THE DISCLOSED EMBODIMENT

FIG. 1 schematically illustrates a gas turbine engine 10 which generally includes a fan section 12, a compressor section 14, a combustor section 16, a turbine section 18, and a nozzle section 20 along a longitudinal axis X. The gas turbine engine 10 of the disclosed embodiment is a relatively low bypass gas turbine engine. Although the disclosed embodiment illustrates a 3-stage fan, a 6-stage compressor, an annular combustor, a single stage high-pressure turbine, and a 2-stage low pressure turbine, various other gas turbine engines will benefit from the present invention.

The engine 10 is configured to provide a variable area turbine nozzle to selectively control the flow of the combustion gas 12 from the combustor section 16 through the turbine section 18. The engine 10 is also referred to as including a Controlled Area Turbine Nozzle (CATN).

Referring to FIG. 2, a turbine nozzle segment 30 includes an arcuate outer vane platform segment 32 and an arcuate inner vane platform segment 34 radially spaced apart from each other. The arcuate outer vane platform segment 32 may form a portion of an outer core engine structure 22 and the arcuate inner vane platform segment 34 may form a portion of an inner core engine structure 24 (FIG. 1) to at least partially define an annular turbine nozzle core gas flow path 26.

The circumferentially adjacent vane platform segments 32, 34 define split lines 36 which thermally uncouple adjacent turbine nozzle segments 30 which may be conventionally sealed therebetween, with, for example only, spline seals. That is, the temperature environment of the turbine section 18 and the substantial aerodynamic and thermal loads are accommodated by the plurality of circumferentially adjoining nozzle segments 30 which collectively form a full, annular ring about the centerline axis X of the engine.

Each turbine nozzle segment 30 includes a multiple (two shown) of circumferentially spaced apart turbine vanes 38, 40 which extend radially between the vane platform segments 32, 34. In the disclosed embodiment, each nozzle segment 30 includes one fixed turbine vane 38 and one rotational turbine vane 40 (doublet) between the vane platform segments 32, 34 to provide a rigid structural assembly which accommodates thermal and aerodynamic loads during operation. That is, the full, annular ring formed by the multiple of turbine nozzle segments 30 provide a vane portion of one stage in the turbine section 18 which is defined by the alternating fixed and rotational turbine vanes 38, 40.

Each turbine nozzle segment 30 includes at least one fixed turbine vane 38 and at least one rotational turbine vane 40 such that the fixed turbine vane 38 and the vane platform segments 32, 34 form a box structure. The vane platform segments 32, 34 may include features 50 to mount each nozzle segment 30 to other engine static structures. It should be understood that although the illustrated embodiment discloses a doublet arrangement, any number of fixed turbine vanes 38 and rotational turbine vanes 40 may be provided in each turbine nozzle segment 30. Movement of the rotational turbine vanes 40 relative the adjacent fixed turbine vanes 38 effectuates a change in throat area formed by the ring of nozzle segments 30 as will be further described below.

Referring to FIG. 3, each turbine vane 38, 40 includes a respective airfoil portion 42F, 42R defined by an outer airfoil wall surface 44F 44R between the leading edge 46F, 46R and a trailing edge 48F, 48R. Each turbine vane 38, 40 may include a fillet 52 to provide a transition between the airfoil portion 42F, 42R and the vane platform segments 32, 34. The outer airfoil wall surface 44 is typically shaped for use in, for example only, a first stage, or other stage, of a high pressure and low pressure stage of the turbine section. The outer airfoil wall 44F, 44R typically have a generally concave shaped portion forming a pressure side 44FP, 44RP and a generally convex shaped portion forming a suction side 44FS, 44RS. It should be understood that respective airfoil portion 42F, 42R defined by the outer airfoil wall surface 44F 44R may be generally equivalent or separately tailored to optimize flow characteristics and transient thermal expansion issues.

An actuator system 54 includes an actuator such as an outer diameter unison ring (illustrated schematically at 56) which rotates an actuator arm 58 and an actuator rod 60 which passes through the inner vane platform segment 32, the rotational turbine vane 40, and the outer vane platform segment 34. The actuator rod 60 rotates each rotational turbine vane 40 about a vane axis of rotation 62 relative the adjacent fixed turbine vanes 38 to selectively vary the turbine nozzle throat area. Since the fixed turbine vane 38 and vane platform segments 32, 34 provide a rigid structure, the rotational turbine vane 40 may include a relatively less complicated rotation, support and sealing structure to provide the variable nozzle throat area capability which minimizes turbine pressure loss, leakage, expense and weight.

The vane axis of rotation 62 is located approximately midway between the trailing edges of an adjacent fixed turbine vanes 38 and rotational turbine vane 40 to close the throat area between the rotational turbine vane 40 and the adjacent fixed turbine vanes 38 on either side of the rotational turbine vane 40 simultaneously (FIG. 4). Airfoils are conventionally rotated around the geometric center of gravity (CG) of the airfoil cross section. Here, the rotational turbine vane 40 vane axis of rotation 62 is biased toward the trailing edge 48R of the rotational turbine vane 40. In one embodiment, a distance L between the trailing edges of an adjacent fixed turbine vanes 38 and rotational turbine vane 40, may be about 1.6 inches (FIG. 4). The rotational turbine vane 40 axis of rotation 62 is then positioned at L/2 or about 0.8 inches from each adjacent fixed turbine vane 38 such that the axis of rotation 62 is located axially aft of the conventional geometric CG.

In operation, rotation of the rotational turbine vanes 40 between a nominal position and a rotated position selectively changes the turbine nozzle throat area as each rotational turbine vane 40 concurrently changes the throat area between itself and both adjacent fixed turbine vanes 38. Since only half the vanes are rotated, the required rotation is less since rotation changes the throat on both sides simultaneously, with less change in the gas exit angle directed to the turbine blades. Furthermore, since only half of the vanes are rotated, the complexity and load requirements of the actuator system 54 are reduced. The alternating rotational-fixed vane arrangement also facilitates a relatively less complicated rotation, support and sealing structure to provide the variable nozzle throat area capability to minimize turbine pressure loss, leakage, expense and weight.

It should be understood that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the device and should not be considered otherwise limiting.

It should be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit from the instant invention.

Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present invention.

The foregoing description is exemplary rather than defined by the limitations within. Many modifications and variations of the present invention are possible in light of the above teachings. The disclosed embodiments of this invention have been disclosed, however, one of ordinary skill in the art would recognize that certain modifications would come within the scope of this invention. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than as specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention. 

1. A turbine nozzle segment for a gas turbine engine comprising: a fixed turbine vane; and a rotational turbine vane adjacent said fixed turbine vane, said rotational turbine vane rotatable about an axis of rotation relative said fixed turbine vane.
 2. The turbine nozzle segment as recited in claim 1, wherein said fixed turbine vane and said rotational turbine vane are located between an outer vane platform segment and an inner vane platform segment.
 3. The turbine nozzle segment as recited in claim 2, wherein said fixed turbine vane is fixed to said outer vane platform segment and said inner vane platform segment.
 4. The turbine nozzle segment as recited in claim 1, further comprising an actuator system having an actuator rod mounted through said rotational turbine vane at said axis of rotation.
 5. The turbine nozzle segment as recited in claim 4, wherein said rotational turbine vane is rotated by a unison ring driven linkage.
 6. The turbine nozzle segment as recited in claim 1, wherein said axis of rotation is aft of a geometric center of gravity of a cross section of said rotational turbine vane.
 7. The turbine nozzle segment as recited in claim 1, wherein said axis of rotation is located approximately midway between a trailing edge of said fixed turbine vane and a trailing edge of said rotational turbine vane.
 8. A turbine section of a gas turbine engine comprising: an annular outer vane platform; an annular inner vane platform; a fixed turbine vane fixed to said annular outer vane platform and said annular inner vane platform; and a rotational turbine vane between said annular outer vane platform and said annular inner vane platform, said rotational turbine vane rotatable about an axis of rotation aft of a geometric center of gravity of a cross section of said rotational turbine vane.
 9. The turbine section as recited in claim 8, wherein said annular outer vane platform and said annular inner vane platform are formed from a respective multiple of outer vane platform segments and a multiple of inner vane platform segments, wherein each of said multiple of outer vane platform segments and said multiple of inner vane platform segments include at least one of said fixed turbine vanes.
 10. The turbine section as recited in claim 9, wherein each of said multiple of outer vane platform segments and said multiple of inner vane platform segments include at least one of said rotational turbine vane between a first fixed turbine vane and a second fixed turbine vane.
 11. The turbine section as recited in claim 8, wherein said fixed turbine vane alternates with said rotational turbine vane.
 12. The turbine section as recited in claim 8, wherein said axis of rotation is located approximately midway between a trailing edge of said fixed turbine vane and a trailing edge of said rotational turbine vane
 13. A method of varying a turbine nozzle throat area of a gas turbine engine comprising the steps of: (A) locating a rotational turbine vane between a first fixed turbine vane and a second fixed turbine vane; and (B) rotating the rotational turbine vane about an axis of rotation to vary a throat area concurrently between the rotational stator vane and both the first fixed turbine vane and the second fixed turbine vane.
 14. A method as recited in claim 13, wherein said step (A) further comprises: (a) alternating each of a multiple of rotational turbine vanes with a multiple of fixed turbine vanes about a turbine section of the gas turbine engine.
 15. A method as recited in claim 14, wherein said step (B) further comprises: (a) rotating a unison ring to simultaneously rotate each of the multiple of rotational turbine vanes.
 16. A method as recited in claim 13, wherein said step (B) further comprises: (a) locating the axis of rotation aft of a geometric center of gravity of a cross section of the rotational turbine vane. 